1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine blade with tip cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, especially an industrial gas turbine, engine, the turbine includes several stages of turbine blades that rotate within a shroud that forms a small gap between the rotating blade tip and the stationary shroud. The hot gas that flows through the turbine blades can also leak through this small gap as the hot gas flows leaks across the blade tip from the pressure side to the suction side. The blade tip region is difficult to properly cool which creates hot spots on section of the blade tip that eventually erode or corrode away. Engine performance and blade tip life can be increased by minimizing the gap so that less hot gas flow leakage occurs, and adequately cool the blade tip section. The blade tips are also subject to rub against the inner surface of the shroud that forms the blade outer air seal (BOAS). Blade tips include one or more tip rails to minimize the gap leakage and surface area of the tip crown that rubs.
High temperature turbine blade tip section heat load is a function of the blade tip leakage flow. A high leakage flow will induce a high heat load onto the blade tip section. Thus, blade tip section sealing and cooling have to be addressed as a single problem. A prior art turbine blade tip design is shown in FIGS. 1-3 and includes a squealer tip rail that extends around the perimeter of the airfoil flush with the airfoil wall to form an inner squealer pocket. The main purpose of incorporating the squealer tip in a blade design is to reduce the blade tip leakage and also to provide for improved rubbing capability for the blade. The narrow tip rail provides for a small surface area to rub up against the inner surface of the shroud that forms the tip gap. Thus, less friction and less heat are developed when the tip rubs.
Traditionally, blade tip cooling is accomplished by drilling holes into the upper extremes of the serpentine coolant passages formed within the body of the blade from both the pressure and suction surfaces near the blade tip edge and the top surface of the squealer cavity. In general, film cooling holes are built in along the airfoil pressure side and suction side tip sections and extend from the leading edge to the trailing edge to provide edge cooling for the blade squealer tip. Also, convective cooling holes also built in along the tip rail at the inner portion of the squealer pocket provide additional cooling for the squealer tip rail. Since the blade tip region is subject to severe secondary flow field, this requires a large number of film cooling holes that requires more cooling flow for cooling the blade tip periphery. FIG. 1 shows the prior art squealer tip cooling arrangement and the secondary hot gas flow migration around the blade tip section. FIG. 2 shows a profile view of the pressure side and FIG. 3 shows the suction side each with tip peripheral cooling holes for the prior art turbine blade of FIG. 1.
The blade squealer tip rail is subject to heating from three exposed side: 1) heat load from the airfoil hot gas side surface of the tip rail, 2) heat load from the top portion of the tip rail, and 3) heat load from the back side of the tip rail. Cooling of the squealer tip rail by means of discharge row of film cooling holes along the blade pressure side and suction peripheral and conduction through the base region of the squealer pocket becomes insufficient. This is primarily due to the combination of squealer pocket geometry and the interaction of hot gas secondary flow-mixing. The effectiveness induced by the pressure film cooling and tip section convective cooling holes become very limited. In addition, a TBC is normally used on an industrial gas turbine (IGT) airfoil for the reduction of blade metal temperature. However, applying the TBC around the blade tip rail without effective backside convective cooling may not reduce the blade tip rail metal temperature. FIG. 4 shows the state of the art turbine blade tip cooling design of the prior art.
One location on the blade tip that is difficult to adequately cool is the tip corner on the pressure or upstream side of the blade. Film cooling holes are used to discharge a layer of film cooling air onto the pressure side wall just below the tip corner as disclosed in U.S. Pat. No. 7,192,250 B2 issued to Boury et al. on Mar. 20, 2007 and entitled HOOLOW ROTOR BLADE FOR THE FUTURE OF A GAS TURBINE ENGINE, which shows a film cooling hole angled upward toward the tip, corner formed with a pressure side tip rail. The layer of film cooling air flows up and over the top or crown of the pressure side tip rail. The layer of cooling air does not touch the tip crown. Also, because of the presence of the pressure side tip rail, the cooling air flows over the forward side of the tip pocket formed between the pressure side tip rail and the suction side tip rail.
U.S. Pat. No. 7,029,235 B2 issued to Liang on Apr. 18, 2006 and entitled COOLING SYSTEM FOR A TIP OF A TURBINE BLADE discloses a blade with pressure side wall having a film cooling hole to discharge cooling air upward toward the blade tip, the blade tip having a flat crown with a tip cooling hole to discharge cooling air out from the tip crown in a direction straight up. The problem with this tip cooling design is that the cooling hole on the tip will push the layer of cooling air from the pressure side of the blade over the tip crown so that no cooling air reaches the tip surface to provide cooling. The only cooling occurring to the tip is convective cooling due to the cooling air passing through the hole formed within the tip itself.
U.S. Pat. No. 7,351,035 B2 issued to Deschamps et al. on Apr. 1, 2008 and entitled HOLLOW ROTOR BLADE FOR THE TURBINE OF A GAS TURBINE ENGINE, THE BLADE BEING FITTED WITH A “BATHTUB” discloses a blade tip with single tip rail on the suction side, a flat tip crown, and a front tip corner angled toward the oncoming hot gas flow. A cooling hole is formed in the tip corner and angled toward the oncoming hot gas flow and a tip floor or crown cooling hole to discharge cooling air up and into the area above the tip floor upstream from the tip rail. The problem with this design is that the pressure side wall of the tip corner is not cooled, and the cooling air discharge from the tip corner cooling hole is directed upward so as not to cool the tip surface immediately downstream from the cooling hole. Also, the tip floor cooling hole discharges cooling air straight up so that the layer of cooling air from the upstream hole is pushed up and away from the top surface of the tip to produce inadequate cooling thereof.
U.S. Pat. No. 6,602,502 B2 issued to Liang on Aug. 5, 2003 and entitled AIRFOIL TIP SQUEALER COOLING CONSTRUCTION discloses a blade with a tip floor that is not flat formed by two short tip rails on the pressure side and suction, side, and the pressure side wall adjacent to the tip corner includes one or two film cooling holes directed to discharge cooling air upward toward the tip corner. In this design, the layer of cooling air from the pressure side wall hole or holes flows up and over the pressure side tip rail without touching the tip surface immediately downstream from the pressure side tip rail. Inadequate cooling of the tip floor at this location occurs in this design.
U.S. Pat. No. 5,282,721 issued to Kildea on Feb. 1, 1994 and entitled PASSIVE CLEARANCE SYSTEM FOR TURBINE BLADES shows an earlier design in which the blade tip corner is pointed and slanted toward the oncoming hot gas flow, and one cooling hole located on the pressure side wall and just below the tip corner to discharge cooling air toward the oncoming hot gas flow. This design is intended to push away the hot gas flow from the gap. In a second embodiment, a second cooling hole discharges cooling, air from the tip crown surface at an angle toward, the oncoming hot gas flow to also block the hot gas from entering the gap. A film layer of cooling air is not developed onto the surface of the tip crown or floor at the pressure side corner.
This problem associated with turbine airfoil tip edge cooling can be minimized by incorporation of anew and effective blade tip cooling geometry design of the present invention into the prior art airfoil tip section cooling design.